(1) Field of the Invention
The invention is related to a rotorcraft with a fuselage and at least one main rotor, said rotorcraft comprising the features of claim 1.
(2) Description of Related Art
A rotorcraft with a single main rotor and a fuselage defining a cabin and a tail boom of the rotorcraft features in operation an inherent aerodynamic nose-down pitching moment about a rotor center of the single main rotor, with the rotor center as reference point of moments. The aerodynamic nose-down pitching moment is an “aerodynamic” load that results from a given main rotor downwash and an airflow, which is due to forward flight in operation and which increases with a corresponding velocity of a given incoming air flow. This aerodynamic load depends on a respective shape of the fuselage and on a vertical distance of a center of aerodynamic effort from the rotor center of the single main rotor.
If no measures are taken to avoid the aerodynamic nose-down pitching moment in operation, an equilibrium of moment around a pitch axis of the rotorcraft would be restored by nose-down pitching of the fuselage, until a resulting offset between a center of gravity of the rotorcraft and a respective current thrust vector and a nose-up pitching moment transferred to the fuselage via a rotor mast of the single main rotor provides counteraction. However, in this case a high moment would arise in the rotor mast in operation, if the single main rotor is embodied as a hingeless or a bearingless main rotor, and would lead to high stress and fatigue problems of the main rotor. If the single main rotor is embodied as an articulated main rotor, a larger nose-down pitching of associated fuselages can occur especially at high flight speeds, but in this case a performance penalty due to higher fuselage drag and passenger discomfort would arise.
In order to avoid the above described drawbacks, the aerodynamic nose-down pitching moment of the rotorcraft is counteracted for maintaining the fuselage in a leveled pitch attitude and to mitigate the fatigue problems with respect to the rotor mast. Therefore, a nose-up pitching moment is generated in operation for compensation of the aerodynamic nose-down pitching moment. This nose-up pitching moment increases with flight speed similarly to the aerodynamic nose-down pitching moment, which also increases with flight speed.
Generation of a nose-up pitching moment is usually achieved by employing a wing which is arranged at a position located at a rear end of the tail boom of the rotorcraft and which is generally designated as a “horizontal stabilizer”. This position allows combining the task of moment generation with the task of stabilizing the pitch motion of the rotorcraft by aerodynamic damping.
If such a horizontal stabilizer is rigidly attached with a predetermined inclination to the fuselage of the rotorcraft, it is hereinafter referred to as a “fixed horizontal stabilizer”. It counters the inherent aerodynamic nose-down pitching moment occurring in operation by generating a respectively required downforce in operation. This downforce is, however, detrimental in terms of performance, as it increases a respectively required rotor thrust, and in terms of stress load on the tail boom. Furthermore, the downforce generated by such a fixed horizontal stabilizer increases with forward flight speed, whereas a given rotor thrust limit decreases due to limited blade loading, particularly on the retreating rotor blades, which ultimately defines a high-speed limit of an associated flight envelope. The fixed horizontal stabilizer is employed by a vast majority of rotorcrafts having a conventional configuration with one single main rotor and an associated tail rotor, having the coaxial rotor configuration or having the intermeshing rotor configuration, which is also known as Flettner rotor system.
Alternatively, a movable horizontal stabilizer can be employed where the inclination of the horizontal stabilizer is continuously controlled by an actuator. Such movable horizontal stabilizers are often used to avoid pitch-up phenomena in low speed forward flight and only allow, as a by-product, to fine-tune the nose-up pitching moment applied to the fuselage for better trim and rotor mast moment relief. Such a movable horizontal stabilizer may be controlled by a mechanical control linkage to an underlying rotor control system. However, the movable horizontal stabilizer requires the actuator and/or the mechanical control linkage to the rotor control system, which increases weight and maintenance effort for the rotorcraft and raises questions about its reliability.
Another possibility to counter the inherent aerodynamic nose-down pitching moment occurring in operation consists in positioning a wing somewhere in front of the rotor center of the rotorcraft. A wing which is positioned somewhere in front of the rotor center of the rotorcraft is generally designated as a “canard wing” and can be adapted to generate a nose-up pitching moment by generating an upward lift force in operation, i.e. a positive lift force.
Furthermore, fixed wings installed directly below the rotor center of the rotorcraft can be designed to provide an upward lift force in operation during forward flight in order to reduce the rotor thrust and, thus, allow for higher flight speeds or improvement of power consumption. Frequently, rotorcrafts with such fixed wings are also equipped with an additional propulsion system to lower the rotor thrust even further.
However, a fixed wing that is positioned directly below the rotor center has only comparatively small influence on pitching moment equilibrium; hence the nose-up pitching moment still needs to be generated by a horizontal stabilizer through down force. Furthermore, a comparatively large download is generated on the fixed wing itself in hover or low forward flight speeds of the rotorcraft due to its position in a corresponding rotor downwash of the main rotor.
The document U.S. Pat. No. 5,454,530 describes a dual-mode high speed rotorcraft with canard wings and a high-lift tail, which includes a rotor for propulsion during low-speed flight and hover. In this rotorcraft, the canard wings and the high-lift tail function together in order to provide substantially all lift for the rotorcraft during the transition between low and high-speed flight, so that the rotor may be unloaded while starting and stopping.
However, any fixed canard wing placed in front of the center of gravity of the rotorcraft has a destabilizing effect on the pitch motion of the rotorcraft, which either prohibits its use or at least severely limits its effect. A corresponding stability margin of the helicopter pitch stability is already crossed at small areas of the canard wing and, thus, severely limits its size and, hence, its beneficial effect. Furthermore, the destabilizing effect has an adverse influence on the handling qualities of the rotorcraft. Finally, a comparatively large download on the fixed canard wing in hover or low forward flight speeds is generated due to its positioning in the rotor downwash. In order to avoid occurrence of these drawbacks, rotatable canard wings can be employed, as described in the following.
The document U.S Pat. Nos. 2007/0095970 A1 discloses an aircraft including an airframe having a fuselage extending between a nose end and a tail end with wings extending laterally from the fuselage. The aircraft includes a rotor that is rotatably mounted on the airframe including a plurality of blades. The wings are described as fixed, but an embodiment is considered where the wings may be rotated between a forward flight position and a vertical flight position, and may also be moved to intermediate flight positions. By rotating the wings into their vertical flight position, download in hover can be minimized, and by rotating them into their forward flight position, control of the pitch motion of the aircraft can be assisted, which then operates in an “airplane mode” with a stopped rotor, so that the wings may generate part of the required lift forces.
However, such rotatable wings need to be controlled by an active control system using e.g. suitable actuators to vary their incidence. But such an active control system has weight and reliability issues. Furthermore, actuation failures of such an active control system may result in instability of the rotorcraft about the pitch axis and lead to a loss of control over the rotorcraft.
Document WO 03/106259 A2 describes an aircraft comprising a boom which has opposite distal end and proximal ands, the latter being pivotally supported on the fuselage.
Document U.S. Pat. No. 2,414,258 A discloses a helicopter having three control panels located within the rotor downwash and mounted radially to the rotor mast on axis around which they can perform relatively free turning movements.
Document U.S. Pat. No. 3,430,894 A also describes a free floating wing for an aircraft, which offers inherent stability when a wind gust contacts the wing—the wing reacts to the hitting current by turning into the direction of the wing sufficiently to equalize the resulting lift.
It should be noted that at least part of the above described systems and devices for generating lift and/or the nose-up pitching moment by wing-type aerodynamic devices in order to counter the inherent aerodynamic nose-down pitching moment occurring in operation of the rotorcraft are also described in the documents U.S. Pat. No. 8,376,264 B1, WO 2010/017397 A1, US 2009/0250548 A1, CA 2 659 499 A1, US 2009/0206208 A1, MC-200107 A, WO 2008/003455 A1, US 2007/0290099 A1, CN 1824576 A, US 2008/0135677 A1, US 2007/0095969 A1, US 2007/0080257 A1, GB 0 619 167 DO, US 2006/0266879 A1, US 2005/0224633 A1, U.S. Pat. No. 6,923,404 B1, US 2004/0108410 A1, US 2004/0093130 A1, U.S. Pat. No. 6,745,979 B1, US 2004/0061025 A1, US 2004/0056144 A1, U.S. Pat. No. 6,669,137 B1, WO 2003/099653 A1 and U.S. Pat. No. 6,622,962 B1. However, all of these systems and devices have at least one of the above described drawbacks.